1. Field of the Invention
The present invention relates to the cooling of gas turbine engine components and, more particularly, to a system for providing tailored pressure boosted cooling flows for high pressure compressor and turbine components.
2. Discussion of the Background Art
Gas turbine engines typically include cooling systems which provide cooling air to turbine rotor components, such as turbine blades, in order to limit the material temperatures experienced by such components. Prior art cooling systems usually acquire the air used to cool turbine components from the engine""s compressor, after which it is diverted and subsequently directed to the turbine section of the engine through an axial passageway. A device commonly known as an inducer is generally employed, at the end of the passageway, to accelerate and direct the airflow tangential to and in the same direction of the rotating rotor. Such inducers, frequently in the form of a circumferentially disposed array of vanes, are used to control the tangential speed and direction of the airflow so that it is substantially equal to that of the turbine rotor. An exemplary inducer utilized for such purpose is disclosed in U.S. Pat. No. 4,882,902 to James R. Reigel et al., entitled xe2x80x9cTurbine Cooling Air Transferring Apparatusxe2x80x9d. Another inducer performing a similar function to the vane-type inducer is disclosed in U.S. Pat. No. 5,245,821 to Theodore T. Thomas Jr. et al. entitled xe2x80x9cStator to Rotor Flow Inducer,xe2x80x9d where a plurality of cylindrical airflow passages are disposed circumferentially about the engine centerline and includes cooling airflow holes or passages that are acutely angled in a tangential manner to the rotational direction of the rotor. The passages include a downstream angled outlet in the form of an open channel that is angled in a rotational direction of the rotor and has a back wall that is at a small acute angle with respect to a plane perpendicular to a centerline of the rotor.
Modern aircraft gas turbine engines are being designed with high overall pressure ratios to increase engine efficiency. Such engines can have compressor discharge or exit temperatures in the regime of 1600xc2x0 F. and higher which can exceed the operating temperature capabilities of component materials. Accordingly, conventional cooling using compressor discharge air is not always feasible because of the high temperature of this air. One technique to reduce the temperature of the compressor discharge air for cooling purposes is cool this air with a cooler fluid, such as fuel, in a heat exchanger as shown in U.S. Pat. No. 5,619,855.
Aircraft gas turbine engine designers constantly strive to improve the efficiency of the gas turbine engine. The use of cooling air increases fuel consumption and, therefore, it is highly desirable to minimize the amount of engine work used to produce the cooling air. The pressure requirements for cooling high pressure compressor components is greater than that of the turbine components but uses a much smaller percentage of cooling airflow, perhaps about 10% of the total cooling airflow. The prior art teaches boosting all of the cooling airflow with an air powered turbo-compressor or other type of supplemental compressor located outside of the engine casing. This wastes fuel by boosting the pressure of the cooling airflow portion that goes to the turbine components to excessive levels. Turbo-compressors are heavy and, therefore, add weight and complexity to the engine. Accordingly, it is highly desirable to have an engine cooling system capable of efficiently supplying high pressure cooling air to high pressure compressor and turbine components of an aircraft gas turbine engine without wasting engine power.
A gas turbine engine cooling system for providing cooling air to engine components includes a core engine having a core flowpath therethrough and, in downstream serial flow relationship, a high pressure compressor, a combustor, and high pressure turbine drivingly connected to the high pressure compressor. A first flowing means is provided for flowing a portion of pressurized air from the high pressure compressor to a heat exchanger to cool the pressurized air and provide cooling air. A second flowing means is provided for flowing a first portion of the cooling air from the heat exchanger to a compressor impeller for boosting pressure of the first portion of the cooling air. The second flowing means is also used to flow a second portion of the cooling air to a turbine cooling means for cooling components of the high pressure turbine. The compressor impeller is operably connected to a compressor disk of the high pressure compressor.
The compressor impeller preferably has a first plurality of compressor radial impeller passages with compressor impeller inlets located on a downstream facing side of a downstream tapering conical shaft section of a high pressure rotor connected to a downstream facing side of the compressor disk. A compressor inducer is operably disposed to channel the first portion of the cooling air into the compressor impeller inlets in a direction substantially tangential to the compressor disk. The heat exchanger may be a fuel to air heat exchanger for cooling the portion of the pressurized air from the first flowing means with fuel. The combustor may be connected to the heat exchanger to receive the fuel from the heat exchanger after the fuel has been used for cooling the portion of the pressurized air from the first flowing means.
One embodiment of the present invention incorporates a first plurality of vanes positioned circumferentially around and extending radially across the core flowpath and axially between the core compressor and the combustor. The second flowing means includes at least one radial vane airflow passage through each of the first plurality of vanes. A hollow area may also be provided within at least some of the vanes effective for receiving the fuel for injection into the core flowpath through apertures such as atomizers positioned across sides of some of the vanes.
The turbine cooling means may include a turbine impeller for boosting pressure of the second portion of the cooling air and the turbine impeller may include a second plurality of radial impeller passages with turbine impeller inlets located on an upstream facing side of the turbine disk, and a second inducer effective for channeling the second portion of the cooling air into the turbine impeller inlets in a direction substantially tangential to the turbine disk.
The present invention has the advantage of being able to tailor the cooling airflows used to cool portions of the high pressure compressor and high pressure turbine to maximize the overall efficiency of the gas turbine engine. The present invention is less costly to build and maintain and less complicated than using an externally mounted turbo-compressor to boost the pressure of the cooling air. The present invention provides apparatus to supply different pressure levels of the cooling flows directed to sections of the high pressure compressor and turbine sections thereby minimizing any ducting or routing losses and unnecessary compression of cooling air directed to the turbine sections. These losses are due to ducting the air outside the engine casing, compressing the air to excessive levels, cooling the air, and then wasting the energy used to over boost the pressure of cooling air needed for the turbine sections which is lower than that of the high pressure compressor. The turbo-compressor also uses more energy and, thus, the apparatus of the present invention is more fuel efficient and less costly to operate because of its ability to tailor the boost pressure of the cooling airflow that are directed to the high pressure compressor and turbine components.